1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine airfoil with near wall cooling.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine, such as an industrial gas turbine (IGT) engine, includes a turbine with a number of stages or rows of turbine stator vanes and rotor blades that react with a hot gas flow. These vanes and blades both include an airfoil portion that is exposed to the hot gas flow. In order to allow for higher temperatures, these airfoils include internal cooling circuits to limit thermal damage to the high gas flow temperature.
Turbine airfoils are typically cast from a ceramic core using the well known investment casting process. However, of the investment casting process has limits as to the size of cooling passages or the shape of cooling channels. Also, thin near wall cooled airfoil surfaces cannot be cast using the investment casting process because the airfoil wall is too thin for casting. The molten liquid metal will not flow through a space formed within the mold that is too thin. Also, small features such as small impingement holes or trip strips cannot be formed using this process because the ceramic core used to form these passages would be much too brittle and small so that the ceramic pieces would break when the heavy and viscous molten metal strikes the pieces.
One prior art cooling design for a turbine airfoil is shown in FIG. 1 and includes an airfoil main body a number of radial channels with re-supply holes in conjunction with film cooling holes. The cooling air is supplied from the fire tree root to the individual radial flow channels and then discharged into the collector cavity formed between the pressure side and suction side walls. The spent cooling air in the collection cavity then flows through the metering holes and into the leading edge impingement cavity to provide impingement cooling for the back side wall of the leading edge. The spent impingement cooling air then flows through the film cooling holes and gill holes as film cooling air for the leading edge surface of the airfoil. With this design, the spanwise and chordwise cooling flow control due to the airfoil external hot gas temperature and pressure variation is difficult to achieve. In addition, single radial channel flow is not the best method of utilizing cooling air and therefore results in a low convective cooling effectiveness.